The present invention relates generally to gas turbine engines, and, more specifically, to active clearance control therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the gases in a high pressure turbine (HPT) which is joined by one drive shaft to the compressor.
In a typical turbofan aircraft engine, a fan is mounted upstream from the compressor and is powered by a low pressure turbine (LPT) mounted downstream of the HPT. In marine and industrial (M & I) applications, the LPT powers an external drive shaft for powering a propulsion system or electrical generator in typical applications.
The compression and combustion cycles introduce energy into the pressurized air, with energy extracted from the combustion gases in the turbine stages. Since the HPT is subject to the hottest combustion gases discharged from the combustor, the various components thereof are typically cooled by bleeding a portion of the pressurized air from the compressor. Any air used for turbine cooling is lost from the combustion cycle and therefore reduces overall efficiency of the engine.
Furthermore, each turbine stage includes a row of turbine rotor blades extending radially outwardly from a supporting rotor disk, with the radially outer tips of the blades being mounted inside a surrounding turbine shroud. The shroud is stationary and supported from a surrounding annular turbine case for maintaining a small radial clearance or gap therebetween.
The turbine blades share a common airfoil profile which is optimized for maximizing the efficiency of energy extraction from the combustion gases. Leakage of the combustion gases at the blade tip gaps further decreases efficiency of the engine.
Accordingly, the radial blade tip clearance is made as small as practical but cannot be too small or undesirable rubbing of the blade tips against the turbine shroud can lead to undesirable damage or shortened component life.
Although the blade tip clearance has an initial magnitude when the engine is cold, the size of the gap or clearance will change as the engine is operated and the various components of the turbine are heated or cooled to different temperatures.
Furthermore, as the engine is operated through various levels of power, the turbine components thermally expand and contract which correspondingly affects the size of the blade tip clearance. Since the turbine blades are directly exposed to the hot combustion gases during operation, they are heated quickly and expand radially outwardly toward the surrounding turbine shroud.
Correspondingly, the turbine shroud is a stationary component supported from the surrounding case and therefore has a different rate of thermal expansion and contraction than the turbine blades mounted on their supporting rotor disk.
The typical turbofan aircraft engine initially operates at a low power, idle mode and then undergoes an increase in power for takeoff and climb operation. Upon reaching cruise at the desired altitude of flight, the engine is operated at lower, or intermediate power setting. The engine is also operated at lower power as the aircraft descends from altitude and lands on the runway, following which thrust reverse operation is typically employed with the engine again operated at high power.
In the various transient modes of operation of the engine where the power increases or decreases, the turbine shroud and blades expand and contract differently, which in turn affects the blade clearance. In one particularly problematic mode of operation called reburst, engine power is quickly increased which correspondingly causes the turbine rotor blades to expand radially outwardly at a greater rate than that of the surrounding turbine shroud. The radial clearance therebetween will therefore decrease during this transient phase.
And, in order to avoid undesirable blade tip rubs against the turbine shroud the initial blade tip clearance must be set sufficiently large, which as indicated above will decrease overall efficiency of the engine due to blade tip leakage.
In order to better control the variable blade tip clearance during engine operation, various clearance control configurations are known, including active clearance control (ACC). In active clearance control, relatively cool fan air or relatively hot compressor bleed air, or a mixture thereof, are channeled to the turbine case from which the shrouds are suspended. The case is either heated or cooled as required to minimize the blade tip clearance specifically during cruise operation of the aircraft where maximum efficiency is desired.
Nevertheless, the effectiveness of conventional active clearance control systems is limited and still requires a relatively large nominal blade tip clearance to avoid undesirable tip rubs, particularly during the reburst condition.
Furthermore, the HPT blades are also typically cooled using a portion of the compressor discharge pressure (CDP) air bled from the last stage of the compressor. The air is suitably channeled through internal cooling channels inside the hollow blades and discharged through the blades in various rows of film cooling holes from the leading edge and aft therefrom, and also typically including a row of trailing edge outlet holes or slots on the airfoil pressure side. This blade cooling air bypasses the combustion process and therefore further reduces efficiency of the engine.
Accordingly, it is desired to provide a gas turbine engine having improved active clearance control and efficiency.